Gas turbine engine components with optimized leading edge geometry

ABSTRACT

A gas turbine engine component is provided. The gas turbine engine component comprises a main body having a leading edge and a leading edge wall including an elongated transition portion extending between the leading edge and a proximate flowpath surface of the main body. The elongated transition portion has an axial length that is greater than a radial height. A gas turbine engine is also provided.

GOVERNMENT LICENSE RIGHTS

This disclosure was made with government support under Contract No.FA8650-09-D-2923 0021 awarded by the United States Air Force. Thegovernment has certain rights in the disclosure.

FIELD

The present disclosure relates to gas turbine engines. In particular,the disclosure relates to gas turbine engine components with optimizedleading edge geometry.

BACKGROUND

Gas turbine engines, and other turbomachines, include multiple sections,such as a fan section, a compressor section, a combustor section, aturbine section, and an exhaust section. Air moves into the enginethrough the fan section. Airfoil arrays in the compressor section rotateto compress the air, which is then mixed with fuel and combusted in thecombustor section. The products of combustion are expanded to rotatablydrive airfoil arrays in the turbine section. Rotating the airfoil arraysin the turbine section drives rotation of the fan and compressorsections.

A blade outer air seal (BOAS) array includes blade outer air seal (BOAS)segments circumferentially disposed about at least a portion of theairfoil arrays. As known, the blade outer air seal environment isexposed to temperature extremes and other harsh environmentalconditions, which may affect the integrity of the blade outer air sealsegments. In addition, high relative movements/displacements between theBOAS segment/array (an exemplary gas turbine engine component) andsurrounding static hardware (e.g., stator vanes) due to the varyingthermal environment in the operational temperature range may, inparticular, expose a leading edge portion of the BOAS to high heatloads, potentially shortening BOAS life and/or compelling additionalcooling flow. The leading edge portion of other gas turbine enginecomponents may also be exposed to high heat loads due to high relativemovements/displacements between the static hardware and the gas turbineengine component, potentially shortening the life of the gas turbineengine component and/or compelling additional cooling flow.

SUMMARY

A gas turbine engine component is provided, according to variousembodiments. The gas turbine engine component comprises a main bodyhaving a leading edge and a leading edge wall including an elongatedtransition portion extending between the leading edge and proximateflowpath surface of the main body. The elongated transition portion hasan axial length that is greater than a radial height.

A gas turbine engine is provided, according to various embodiments. Thegas turbine engine comprises a blade stage and a circumferential arrayof blade outer air seal segments in the blade stage. A blade outer sealsegment (BOAS) comprises a main body that extends axially with respectto a central axis from a leading edge portion of the main body to atrailing edge portion of the main body. The leading edge portion of theBOAS includes a leading edge and a leading edge wall including anelongated transition portion extending between the leading edge and aninner diameter flowpath surface of the main body. The elongatedtransition portion has an axial length that is greater than a radialheight.

A gas turbine engine is provided, according to various embodiments. Thegas turbine engine comprises an engine case, a turbine stage comprisinga stator vane and a rotor blade, and a gas turbine engine component. Thegas turbine engine component comprises a main body having a leading edgeand a leading edge wall including an elongated transition portionextending between the leading edge and a proximate flowpath surface ofthe main body. The elongated transition portion has an axial length thatis greater than a radial height.

In any of the foregoing embodiments, a static structure is configured tobe disposed adjacent and upstream of the gas turbine engine component ina gas turbine engine and each of the static structure and the gasturbine engine component is configured to move relative to each otherbecause of thermal or mechanical deflections. The elongated transitionportion has an axial length that is greater than a radial height by upto one order of magnitude. The axial length of the elongated transitionportion is about three to about ten times the radial height of theelongated transition portion. The elongated transition portion isconfigured as an ellipse with an elliptical factor of greater than about3, wherein the elliptical factor is defined as a length of a major axisdivided by the length of a minor axis. The elongated transition portionhas a first tangency point and a second tangency point and the axiallength comprises a length between the leading edge and the secondtangency point. The elongated transition portion has a chamfer of lessthan about 18 degrees combined with a radius. The elongated transitionportion is configured as a chamfer blended with a radius to at least oneof the leading edge or the proximate flowpath surface of the main body.The gas turbine engine component comprises a blade outer air seal (BOAS)segment.

BRIEF DESCRIPTION OF THE DRAWINGS

The drawings described herein are for illustration purposes only and arenot intended to limit the scope of the present disclosure in any way.The present disclosure will become more fully understood from thedetailed description and the accompanying drawings wherein:

FIG. 1 shows a cross-section of a gas turbine engine, according tovarious embodiments;

FIG. 2 schematically shows one of the turbine stages in the turbinesection of the gas turbine engine of FIG. 1 (e.g., a first stage of ahigh pressure turbine (HPT)) and its associated array of blade outer airseal segments (a single blade outer air seal segment is shown),according to various embodiments;

FIG. 3 shows a sectional view of another BOAS segment in isolation, theBOAS segment having an optimized leading edge geometry defined by aleading edge wall including an elongated transition portion extendingbetween a leading edge of the BOAS segment and an inner diameterflowpath surface of the BOAS segment, according to various embodiments;

FIG. 4 is a schematic view of an elongated transition portion of anexemplary BOAS segment having an optimized leading edge geometry, theelongated transition portion configured as an ellipse, according tovarious embodiments; and

FIGS. 5A through 5J are schematic views of the leading edge portion ofan exemplary gas turbine engine component (a BOAS segment in thedepicted embodiment) deflected relative to an upstream static structure(a stator vane in the depicted embodiment), a leading edge wall of theexemplary gas turbine engine component including a conventional(including an untreated leading edge portion) transition portion (FIGS.5A-5C and FIGS. 5E-5F) and an elongated transition portion according tovarious embodiments (FIGS. 5D and 5G-5J).

DETAILED DESCRIPTION

The detailed description of exemplary embodiments herein makes referenceto the accompanying drawings, which show exemplary embodiments by way ofillustration and its best mode, and not of limitation. While theseexemplary embodiments are described in sufficient detail to enable thoseskilled in the art to practice the invention, it should be understoodthat other embodiments may be realized and that logical, chemical andmechanical changes may be made without departing from the spirit andscope of the invention. For example, the steps recited in any of themethod or process descriptions may be executed in any order and are notnecessarily limited to the order presented. Moreover, many of thefunctions or steps may be outsourced to or performed by one or morethird parties. Furthermore, any reference to singular includes pluralembodiments, and any reference to more than one component or step mayinclude a singular embodiment or step. Also, any reference to attached,fixed, connected or the like may include permanent, removable,temporary, partial, full and/or any other possible attachment option.Additionally, any reference to without contact (or similar phrases) mayalso include reduced contact or minimal contact. It is to be understoodthat unless specifically stated otherwise, references to “a,” “an,”and/or “the” may include one or more than one and that reference to anitem in the singular may also include the item in the plural.

Various embodiments are directed to gas turbine engines and gas turbineengine components such as blade outer air seal (BOAS) segments withoptimized leading edge geometry. Relative movement or shifts due to thevarying thermal environment between a non-rotating component and anadjacent BOAS in a turbine or compressor stage of a gas turbine enginecan result in a leading edge portion of the BOAS projecting into the hotcore flowpath of the gas turbine engine, resulting in a high heat loadfor the BOAS leading edge portion, thereby shortening BOAS life and/orcompelling additional cooling. Various embodiments permit the hot coreflowpath air to impinge on the BOAS leading edge portion at a reducedincidence angle (relative to conventional leading edge geometry),thereby minimizing exposure of the BOAS leading edge portion to highheat transfer coefficients from the hot core flowpath air and thusextending BOAS life and/or minimizing cooling requirements. While a BOASsegment having a leading edge portion with an optimized geometry isdescribed herein, it is to be understood that the BOAS segment is anexemplary gas turbine engine component and that other gas turbine enginecomponents may benefit from an optimized leading edge geometry accordingto various embodiments.

According to various embodiments, and with reference to FIG. 1, a gasturbine engine 20 is schematically illustrated. According to variousembodiments, gas turbine engine 20 may be a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28, for example. Alternativeengines might include an augmentor section (not shown) among othersystems or features, according to various embodiments. According tovarious embodiments, the fan section 22 drives air along a bypassflowpath B while the compressor section 24 drives air along a coreflowpath for compression and communication into the combustor section 26then expansion through the turbine section 28. Although depicted as aturbofan gas turbine engine in the disclosed non-limiting embodiment, itshould be understood that the concepts described herein are not limitedto use with turbofans as the teachings may be applied to other types ofturbine engines including three-spool architectures, non-geared turbineengines, and land-based turbines, according to various embodiments.

According to various embodiments, gas turbine engine 20 may generallyinclude a first spool 30 and a second spool 32 mounted for rotationabout an engine central axis A relative to an engine static structure 36via several bearing systems 38. It should be understood that variousbearing systems 38 at various locations may alternatively oradditionally be provided, according to various embodiments. According tovarious embodiments, the first spool 30 may generally include a firstshaft 40 that interconnects a fan 42, a first compressor 43 and a firstturbine 46. According to various embodiments, the first shaft 40 may beconnected to the fan 42 through a gear assembly of a fan drive gearsystem 48 to drive the fan 42 at a lower speed than the first spool 30.According to various embodiments, the second spool 32 may include asecond shaft 49 that interconnects a second compressor 52 and secondturbine 55. According to various embodiments, the first spool 30 may runat a relatively lower pressure than the second spool 32. It is to beunderstood that “low pressure” and “high pressure” or variations thereofas used herein are relative terms indicating that the high pressure isgreater than the low pressure. According to various embodiments, anannular combustor 57 may be arranged between the second compressor 52and the second turbine 55. According to various embodiments, the firstshaft 40 and the second shaft 49 may be concentric and rotate viabearing systems 38 about the engine central axis A which is collinearwith their longitudinal axes, according to various embodiments.

According to various embodiments, the core airflow may be compressed bythe first compressor 43 then the second compressor 52, mixed and burnedwith fuel in the annular combustor 57, then expanded over the secondturbine 55 and first turbine 46. According to various embodiments, thefirst turbine 46 and the second turbine 55 may rotationally drive,respectively, the first spool 30 and the second spool 32 in response tothe expansion. According to various embodiments, gas turbine engine 20may be a high-bypass geared aircraft engine that has a bypass ratio thatis greater than about six (6), with an example embodiment being greaterthan ten (10). According to various embodiments, the gear assembly ofthe fan drive gear system 48 may be an epicyclic gear train, such as aplanetary gear system or other gear system, with a gear reduction ratioof greater than about 2.3:1 and the first turbine 46 may have a pressureratio that is greater than about 5, for example. According to variousembodiments, the first turbine 46 pressure ratio is pressure measuredprior to inlet of first turbine 46 as related to the pressure at theoutlet of the first turbine 46 prior to an exhaust nozzle. According tovarious embodiments, first turbine 46 may have a maximum rotor diameterand the fan 42 may have a fan diameter such that a ratio of the maximumrotor diameter divided by the fan diameter is less than 0.6. It shouldbe understood, however, that the above parameters are only exemplary.

A significant amount of thrust may be provided by the bypass flow B dueto the high bypass ratio. According to various embodiments, the fansection 22 of the gas turbine engine 20 may be designed for a particularflight condition—typically cruise at an airspeed of 0.8 Mach andaltitude of 35,000 feet (10.67 km). The flight condition of 0.8 Mach and35,000 feet (10.67 km) may be a condition at which an engine isoperating at its best fuel consumption. To make an accurate comparisonof fuel consumption between engines, fuel consumption is reduced to acommon metric which is applicable to all types and sizes of turbojetsand turbofans. The term that may be used to compare fuel consumptionbetween engines is thrust specific fuel consumption, or TSFC. This is anengine's fuel consumption in pounds per hour divided by the net thrust.Stated another way, TSFC is the amount of fuel required to produce onepound of thrust. The TSFC unit is pounds per hour per pounds of thrust(lb/hr/lb Fn). When the reference is to a turbojet or turbofan engine,TSFC is often simply called specific fuel consumption, or SFC. “Low fanpressure ratio” is the pressure ratio across the fan blade alone,without a fan exit guide vane system. The low fan pressure ratio asdisclosed herein according to one non-limiting embodiment is less thanabout 1.45. “Low corrected fan tip speed” is the actual fan tip speed infeet per second divided by an industry standard temperature correctionof [(T_(ram)° R)/(518.7° R)]^0.5. The “Low corrected fan tip speed” asdisclosed herein according to one non-limiting embodiment may be lessthan about 1150 feet per second (350 m/s).

Each of the first and second compressors 43 and 52 and first and secondturbines 46 and 55 in the gas turbine engine 20 comprises interspersedstages of rotor blades 70 and stator vanes 72. The rotor blades 70rotate about the centerline with the associated shaft while the statorvanes 72 remain stationary about the centerline. The first and secondcompressors 43 and 52 in the gas turbine engine may each comprise one ormore compressor stages. The first and second turbines 46 and 55 in thegas turbine engine 20 may each comprise one or more turbine stages. Eachcompressor stage and turbine stage may comprise multiple sets ofrotating blades (“rotor blades”) and stationary vanes (“stator vanes”).For example, FIG. 2 schematically shows, by example, a first turbinestage of the second turbine 55 (a high-pressure turbine (HPT)) in theturbine section of the gas turbine engine. Unless otherwise indicated,the term “blade stage” refers to at least one of a turbine stage or acompressor stage.

With continued reference to FIGS. 1 and 2, according to variousembodiments, the depicted first turbine stage of the HPT comprises therotor blade 70 and the stator vane 72. The stator vane 72 may have aninner stator vane platform 63 and an outer stator vane platform 65. Therotor blade 70 may comprise a blade airfoil section 100 and a platform200, such as a rotor blade platform. A blade outer air seal (BOAS)segment 50 is attached to an engine case structure 44 of the gas turbineengine 20 by a receiving portion 68 of the engine case structure 44. TheBOAS segment 50 faces the rotor blade 70 (an exemplary turbine blade inFIG. 2) to define a radial tip clearance 53 between the rotor blade 70and the BOAS segment 50 (more particularly, between a proximate innerdiameter flowpath surface 64 of the BOAS and the turbine blade tip). Aminimal radial tip clearance 53 is sought to be maintained as thesmaller the clearance, the greater the turbine efficiency. The BOASsegment 50 locally bounds the radially outboard extreme of the coreflowpath through the gas turbine engine 20. Although only one BOASsegment 50 is shown in FIG. 2, the turbine stage comprises an associatedarray of blade outer air seal segments. A number of BOAS segments 50 maybe arranged circumferentially about engine axis A to form a shroud,according to various embodiments. According to various embodiments, theBOAS segments 50 may alternatively be formed as a unitary BOASstructure, with the same features described herein.

Still referring to FIG. 2 and now specifically to FIG. 3, according tovarious embodiments, the BOAS segment 50 may include a main body 54 thatextends generally axially from a leading edge portion 56 to a trailingedge portion 58 and from a radially outward facing surface 62 at anoutboard side of BOAS segment 50 to the inner diameter flowpath surface64 at an inboard side of BOAS segment 50. The leading edge portion 56 ofBOAS segment 50 includes a leading edge 76 and a leading edge wall 56 a.In accordance with various embodiments, the leading edge wall 56 aincludes an elongated transition portion 56 b that extends between theleading edge 76 and the inner diameter flowpath surface 64 of the BOASsegment as hereinafter described. The BOAS segment 50 also includes atleast one leading attachment portion 60 a (also referred to as“attachment portions 60 a”) disposed at or near the leading edge portion56 and at least one trailing attachment portion 60 b (also referred toas “attachment portions 60 b”) disposed at or near the trailing edgeportion 58. Each of the attachment portions 60 a, 60 b may define aflange 66. Flange 66 of attachment portions 60 a and/or 60 b may extendin an axially aft direction. Flange 66 of attachment portions 60 aand/or 60 b may alternatively extend in an axially forward direction, asshown in the figures. Flange 66 of attachment portions 60 a and/or 60 bmay alternatively extend in and/or out of the page. Each axiallyextending flange 66 corresponds to the receiving portion 68 of theengine case structure 44 to support and attach the BOAS segment 50(shown schematically in FIG. 2). According to various embodiments, theattachment portions 60 a may be circumferentially offset,circumferentially aligned, or a combination of both, from the attachmentportions 60 b in response to BOAS segment 50 parameters.

Still referring to FIGS. 2 and 3 and now to FIG. 4, according to variousembodiments, and as noted previously, the leading edge portion 56 of theBOAS segment has an elongated transition portion 56 b extending from theleading edge 76 to the inner diameter flowpath surface 64. The elongatedtransition portion 56 b, which may have various configurations ashereinafter described, defines an optimized leading edge geometry thatdeviates away from a conventional leading edge geometry (see, e.g.,FIGS. 5A through 5H). The elongated transition portion 56 b has an axiallength (L) that is greater than a radial height (H), of up to one orderof magnitude. The axial length of the elongated transition portion 56 bis about three to about ten times the radial height of the elongatedtransition portion 56 b. In various embodiments, the axial length may bedefined as the length between a first tangency point 74 or the leadingedge (face) 76 and the leading edge wall 56 a and a second tangencypoint 78 and the inner diameter flowpath surface 64 and the radialheight is defined as the radial distance between the inner diameterflowpath surface 64 and the leading edge wall.

Still referring to FIGS. 2 through 4 and now specifically to FIG. 5D,according to various embodiments, the elongated transition portion 56 bmay be configured as an ellipse. The ellipse may have an ellipticalfactor of greater than about 3, wherein the elliptical factor is definedas a length of a major axis divided by the length of a minor axis.Referring now specifically to FIGS. 5G and 5H, according to variousembodiments, the elongated transition portion may be configured with theellipse (FIG. 5G) or with a chamfer in combination with a radius (FIG.5H). A degree of the chamfer is less than about 18 degrees. The shape ofthe elongated transition portion (the ellipse) in FIG. 5G is the same asthe shape of the elongated transition portion in FIG. 5D.

As herein described, the leading edge portion (more particularly, theelongated transition portion) of the BOAS segment (and other gas turbineengine components) has a geometry such that over an operationaltemperature range, thermal and/or mechanical deflections of anon-rotating structure (e.g., the upstream stator vane 72 depicted inFIG. 2) upstream of the BOAS array relative to thermal and/or mechanicaldeflections of the BOAS array cause relative movement of thenon-rotating structure and the BOAS array to expose the leading edgeportion of the BOAS to hot core flowpath air. The BOAS array may beradially deflected outboard of the upstream non-rotating structure asshown in FIGS. 5A through 5D (a “waterfall condition”) or radiallydeflected inboard of the upstream non-rotating structure as shown inFIGS. 5E through 5J (a “dam condition”).

More specifically, FIG. 5A depicts a BOAS segment/array with anuntreated (i.e., not configured in accordance with various embodiments)leading edge portion. FIGS. 5B and 5C depict a BOAS segment/array with aconventional transition between the leading edge wall and the innerdiameter flowpath surface 64, with FIG. 5B depicting a conventionalchamfer and FIG. 5C depicting a conventional radius. As a result, theleading edge portion of the BOAS segment/array in each of FIGS. 5A, 5B,and 5C is fully exposed to hot core flowpath air (“hot gas flow”). Bycontrast, the BOAS segment/array in FIG. 5D has the elongated transitionportion configured as the ellipse in accordance with variousembodiments, such that the hot core flowpath air is transitioned fromthe upstream vane to the BOAS array with a reduced incidence angle(relative to the conventional or untreated leading edge geometry)accommodating an increased range of relative radial deflection.

As noted previously, the BOAS array may alternatively be radiallydeflected inboard of the upstream non-rotating structure (a “damcondition) as shown in FIGS. 5E through 5H. FIG. 5E depicts a BOASsegment/array with an untreated (i.e., not configured in accordance withvarious embodiments) leading edge portion. FIG. 5F depicts a BOASsegment/array with a conventional chamfered transition between theleading edge wall and the inner flowstream path. As a result of therelative radial deflection between the upstream non-rotating component(a stator vane in the depicted embodiment) and the adjacent BOASsegment/array, the leading edge portion of the BOAS segment/array ineach of FIGS. 5E and 5F is fully exposed to hot core flowpath air andhigh heat transfer coefficients. By contrast, the BOAS segment/array inFIG. 5G has the elongated transition portion configured as an ellipse inaccordance with various embodiments, so as to permit more deflection ofthe hot core flowpath air off the leading edge portion of the BOASsegment/array relative to an untreated or conventional BOASsegment/array (FIGS. 5A-5C and 5E-5F). As noted previously, the BOASsegment/array in FIG. 5H has an elongated transition portion configuredas a chamfer with a blended radius in accordance with variousembodiments, so as to also permit more deflection of the hot coreflowpath air off the leading edge portion of the BOAS segment/arrayrelative to an untreated or conventional BOAS segment/array (FIGS. 5A-5Cand 5E-5F). The elongated transition portion configured as a chamferwith blended radii as depicted in FIG. 5H is also easier to make andinspect relative to other transitions. FIG. 5I is a schematic view ofthe leading edge portion of an exemplary BOAS segment in the depictedembodiment) illustrating that a transition of the elongated transitionportion to the leading edge (face) of the BOAS segment may produce acorner rather than a tangent as in FIGS. 5D and 5G-5H, according tovarious embodiments. FIG. 5J is a schematic view of the leading edgeportion of an exemplary BOAS segment illustrating an elongatedtransition portion comprising a short chamfer between two radii.

During gas turbine engine 20 operation, and over the operationaltemperature range, the BOAS segment 50 is subjected to different thermalloads and environmental conditions (i.e., the thermal environmentsurrounding each turbine or compressor stage varies during operation).As a result, the thermal and/or mechanical deflections of thenon-rotating structure adjacent to the BOAS segment array and thethermal and/or mechanical deflections of the BOAS segment array may besuch that relative movement exposes the leading edge portion to hot coreflowpath air. According to various embodiments, the leading edge portionis configured such that the hot core flowpath air is transitioned fromthe upstream non-rotating structure (e.g., the upstream stator vane) tothe BOAS array with a reduced incidence angle that accommodates anincreased range of relative radial deflections. The variation in theradial clearance between the stationary vane and the adjacent BOAS is aresult of how the outer stator vane platform 65 and the engine casestructure 44 react different to the varying thermal environment.

For example, in the first stage of the high pressure turbine (HPT) (thedesignation “T1” referring to the first stage of the HPT) depicted inFIG. 2, a transition of the outer stator vane platform 65 to the leadingedge of the BOAS is defined to yield a smooth core flowpath at steadystate conditions for maximum efficiency, typically at cruise conditions.However, relative motion or shifts between the T1 stator vane and the T1BOAS segment due to the varying thermal environment can result in theBOAS leading edge portion moving radially outboard of the outer statorvane platform 65 (outer diameter platform) of the T1 stator vane,creating an outward step in the hot core flowpath from the outer statorvane platform 65 and BOAS (see, e.g., FIGS. 5A through 5D). Relativemotion or shifts may also result in the BOAS leading edge portion movingradially inboard of the outer stator vane platform 65 (outer diameterplatform) of the T1 stator vane, making the BOAS leading edge portionproject into the hot core flowpath (see, e.g., FIGS. 5E through 5H). Asnoted previously, this results in higher heat load for the BOAS leadingedge portion, shortening its life and/or compelling additional cooling.

However, according to various embodiments, the elongated transitionportion 56 b may improve gas flow transition across the leading edgewall 56 a, and may prevent a stagnation region at the leading edgeportion 56. More particularly, various embodiments permit the transitionfrom the upstream stator vane to the leading edge of the BOAS to besmoother and the leading edge portion less sensitive to being projectedinto the hot core flowpath air as a result of the relativemovement/shifting of the BOAS segment and the surrounding staticstructure (e.g., the upstream stator vanes). As a result, variousembodiments prolong BOAS life and/or tend to minimize cooling flowrequirements for the BOAS segment/array, thereby maximizing turbineefficiency.

While various embodiments have been described to ease the transitionbetween an upstream stator vane and an adjacent BOAS segment in aturbine stage, it is to be understood that various embodiments may beused to smooth the transition between adjacent non-rotating structures.As depicted in FIG. 2, the proximate flowpath surface for the BOASleading edge 76 of the main body 54 of BOAS segment 50 is the innerdiameter flowpath surface 64. Similar to elongated transition portion 56b of BOAS segment 50, the stator vane 72 has an elongated transitionportion 65 b on outer stator vane platform 65 that transitions from astatic combustor panel (not shown) at an outer platform leading edge 77to the proximate flowpath surface 65 c, the boundary between outerstator vane platform 65 and an airfoil 71 of stator vane 72. Likewise,inner stator vane platform 63 includes an elongated transition portion63 b that extends from an inner platform leading edge 79 to a proximateflowpath surface 63 c, the boundary between the elongated transitionportion 63 b and the airfoil 71 of stator vane 72. Hence, the “proximateflowpath surface” may be an inner flowpath surface or an outer flowpathsurface.

In addition, while the first turbine stage of a HPT is depicted in FIG.2, it is to be understood that various embodiments may be utilized forstatic gas turbine engine components in any turbine stage of the HPT orlow-pressure turbine (LPT) (i.e., first turbine) and in any compressorstage of the high-pressure compressor (HPC) (i.e., second compressor) orthe low-pressure compressor (LPC) (i.e., first compressor). While a BOASsegment having specially configured leading edge geometry for deflectinghot core flowpath air has been described in accordance with variousembodiments, it is to be understood that other gas turbine enginecomponents may benefit from an optimized leading edge geometry accordingto various embodiments. For example, mechanical and thermal deflectionsof a non-rotating structure adjacent to a gas turbine engine componentmay be such that relative movement exposes a leading edge portion of thegas turbine engine component to hot core flowpath air. The leading edgeportion of the gas turbine engine component may be configured with theoptimized leading edge geometry in accordance with various embodimentssuch that the hot core flowpath air is transitioned from the upstreamnon-rotating structure to the gas turbine engine component with reducedincidence angle that accommodates an increased range of relative radialdeflections. Other exemplary gas turbine engine components that maybenefit from various embodiments include, but are not limited to,combustor panels, vane platforms, BOAS, Mid-turbine frames, Transitionducts, etc. In the detailed description herein, references to “variousembodiments”, “one embodiment”, “an embodiment”, “an exampleembodiment”, etc., indicate that the embodiment described may include aparticular feature, structure, or characteristic, but every embodimentmay not necessarily include the particular feature, structure, orcharacteristic. Moreover, such phrases are not necessarily referring tothe same embodiment. Further, when a particular feature, structure, orcharacteristic is described in connection with an embodiment, it issubmitted that it is within the knowledge of one skilled in the art toaffect such feature, structure, or characteristic in connection withother embodiments whether or not explicitly described. After reading thedescription, it will be apparent to one skilled in the relevant art(s)how to implement the disclosure in alternative embodiments.

Benefits, other advantages, and solutions to problems have beendescribed herein with regard to specific embodiments. However, thebenefits, advantages, solutions to problems, and any elements that maycause any benefit, advantage, or solution to occur or become morepronounced are not to be construed as critical, required, or essentialfeatures or elements of the invention. The scope of the invention isaccordingly to be limited by nothing other than the appended claims, inwhich reference to an element in the singular is not intended to mean“one and only one” unless explicitly so stated, but rather “one ormore.” Moreover, where a phrase similar to “at least one of A, B, or C”is used in the claims, it is intended that the phrase be interpreted tomean that A alone may be present in an embodiment, B alone may bepresent in an embodiment, C alone may be present in an embodiment, orthat any combination of the elements A, B and C may be present in asingle embodiment; for example, A and B, A and C, B and C, or A and Band C. Furthermore, no element, component, or method step in the presentdisclosure is intended to be dedicated to the public regardless ofwhether the element, component, or method step is explicitly recited inthe claims. No claim element herein is to be construed under theprovisions of 35 U.S.C. 112(f) unless the element is expressly recitedusing the phrase “means for.” As used herein, the terms “comprises”,“comprising”, or any other variation thereof, are intended to cover anon-exclusive inclusion, such that a process, method, article, orapparatus that comprises a list of elements does not include only thoseelements but may include other elements not expressly listed or inherentto such process, method, article, or apparatus.

What is claimed is:
 1. A gas turbine engine component comprising: a mainbody having a leading edge and a leading edge wall including anelongated transition portion configured as a section of an ellipseextending between the leading edge and a proximate flowpath surface ofthe main body, wherein the section comprises a portion defined betweenan intersection of a major axis with the ellipse and an intersection ofa minor axis with the ellipse, wherein the length of the major axis isgreater than the length of the minor axis, and wherein the elongatedtransition portion has an axial length that is greater than a radialheight.
 2. The gas turbine engine component of claim 1, wherein a staticstructure is configured to be disposed adjacent and upstream of the gasturbine engine component in a gas turbine engine and each of the staticstructure and the gas turbine engine component is configured to moverelative to each other because of thermal or mechanical deflections. 3.The gas turbine engine component of claim 2, wherein the elongatedtransition portion has an axial length that is greater than a radialheight by up to one order of magnitude.
 4. The gas turbine enginecomponent of claim 3, wherein the axial length of the elongatedtransition portion is about three to about ten times the radial heightof the elongated transition portion.
 5. The gas turbine engine componentof claim 3, wherein the section of an ellipse has an elliptical factorof greater than about 3, wherein the elliptical factor is defined as alength of the major axis divided by the length of the minor axis,wherein the major axis is parallel to an axis of rotation of the gasturbine engine.
 6. The gas turbine engine component of claim 5, whereinthe elongated transition portion has a first tangency point and a secondtangency point and the axial length comprises a length between the firsttangency point and the second tangency point.
 7. The gas turbine enginecomponent of claim 3, wherein the elongated transition portion has achamfer of less than about 18 degrees combined with a radius.
 8. The gasturbine engine component of claim 7, wherein the elongated transitionportion is configured as a chamfer blended with a radius to at least oneof the leading edge or the proximate flowpath surface of the main body.9. The gas turbine engine component of claim 1, wherein the gas turbineengine component comprises a blade outer air seal (BOAS) segment.
 10. Agas turbine engine comprising: a blade stage; and a circumferentialarray of blade outer air seal segments in the blade stage, a blade outerseal (BOAS) segment comprising: a main body that extends axially withrespect to a central axis from a leading edge portion of the main bodyto a trailing edge portion of the main body, wherein the leading edgeportion of the BOAS segment includes a leading edge and a leading edgewall including an elongated transition portion extending between theleading edge and an inner diameter flowpath surface of the main body,wherein the elongated transition portion has an axial length that isgreater than a radial height, wherein the elongated transition portionis configured as a section of an ellipse, wherein the section comprisesa portion defined between an intersection of a major axis with theellipse and an intersection of a minor axis with the ellipse, andwherein the length of the major axis is greater than the length of theminor axis.
 11. The gas turbine engine of claim 10, wherein the axiallength of the elongated transition portion is about three to about tentimes the radial height of the elongated transition portion.
 12. The gasturbine engine of claim 10, wherein the ellipse has an elliptical factorof greater than about 3, wherein the elliptical factor is defined as alength of the major axis divided by the length of the minor axis,wherein the major axis is parallel to an axis of rotation of the gasturbine engine.
 13. The gas turbine engine of claim 10, wherein theelongated transition portion has a chamfer of less than about 18degrees.
 14. The gas turbine engine of claim 13, wherein the elongatedtransition portion is configured as a chamfer blended with a radius toat least one of the leading edge or the inner diameter flowpath surfaceof the main body.
 15. A gas turbine engine comprising: an engine case; aturbine stage comprising a stator vane and a rotor blade; and a gasturbine engine component comprising a main body having a leading edgeand a leading edge wall including an elongated transition portionconfigured as a section of an ellipse extending between the leading edgeand proximate flowpath surface of the main body, wherein the sectioncomprises a portion defined between an intersection of a major axis withthe ellipse and an intersection of a minor axis with the ellipse,wherein the length of the major axis is greater than the length of theminor axis, and wherein the elongated transition portion has an axiallength that is greater than a radial height.
 16. The gas turbine engineof claim 15, wherein the axial length of the elongated transitionportion is about three to about ten times the radial height of theelongated transition portion.
 17. The gas turbine engine of claim 15,wherein the elongated transition portion is further configured as atleast one of a chamfer of less than 18 degrees or a chamfer of less than18 degrees blended with a radius to at least one of the leading edge orthe proximate flowpath surface of the main body.
 18. The gas turbineengine of claim 17, wherein the elongated transition portion isconfigured as at least one of the ellipse having an elliptical factor ofgreater than about 3 wherein the major axis is parallel to an axis ofrotation of the gas turbine engine, a chamfer of less than about 18degrees, or a chamfer blended with a radius to at least one of theleading edge or the proximate flowpath surface of the main body.
 19. Thegas turbine engine of claim 15, wherein the gas turbine engine componentcomprises a blade outer air seal segment attached to the engine case andfacing the rotor blade to locally bind a radially outboard extreme of acore flowpath through the gas turbine engine, the blade outer air sealsegment comprising: a main body that extends axially with respect to acentral axis from a leading edge portion of the main body to a trailingedge portion of the main body, wherein the leading edge portion includesa leading edge wall; and an elongated transition portion included in theleading edge wall, the elongated transition portion having an axiallength that is greater than a radial height of the elongated transitionportion by up to one order of magnitude.
 20. The gas turbine engine ofclaim 15, wherein a static structure is configured to be disposedadjacent and upstream of the gas turbine engine component in the gasturbine engine and each of the static structure and the gas turbineengine component is configured to move relative to each other because ofthermal or mechanical deflections.